Coolant flow redirection component

ABSTRACT

A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section, and a turbine section fluidly connected to the combustor and mechanically connected to the compressor section via a shaft. Multiple rotors are disposed in one of the compressor section and the turbine section. Each of the rotors includes a rotor disk portion having a radially inward bore, and is static relative to the shaft. Each rotor is axially adjacent at least one other rotor and a gap is defined between each rotor and an adjacent rotor. A cooling passage for a cooling flow is defined between the shaft and the rotors, and a cooling flow redirection component is disposed at the gap and is operable to redirect the cooling flow in the cooling passage into the gap.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.14/841,265 filed Aug. 31, 2015. U.S. patent application Ser. No.14/841,265 claims priority to U.S. Provisional Application No.62/045,749 filed on Sep. 4, 2014.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This disclosure was made with Government support under FA8650-09-D-2923awarded by The United States Air Force. The Government has certainrights in this invention.

TECHNICAL FIELD

The present disclosure relates generally to rotor disk configurations ina gas turbine engine, and more specifically to rotor disk cooling.

BACKGROUND

Gas turbine engines, such as those utilized in military and commercialaircraft utilize a compressor portion to compress air, a combustorportion to mix the compressed air with a fuel and ignite the mixture,and a turbine portion to expand the resultant gasses from the ignition.The expansion of the gasses in the turbine section drives the turbinesection to rotate. The turbine section is tied to the compressor sectionby at least one shaft, and the rotation of the turbine section drivesthe rotation of the compressor section.

Included in the compressor section and the turbine section are multiplerotors. Each of the rotors includes a radially inward rotor disk andmultiple rotor blades protruding radially outward from the rotor disk.The rotor blades are airfoil blades that protrude into the flow pathdefined by the compressor section and the turbine section. The rotordisks, and particularly the radially inward portions of the rotor disks,are thicker and take longer to heat up and cool down than the radiallyoutward portion of the disks and the blades extending from the disks. Asa result, thermal gradients can occur across the rotor disk. The thermalgradients reduce the effective lifespan of the disk.

SUMMARY OF THE INVENTION

In one example embodiment of the following disclosure, a gas turbineengine includes a compressor section, a combustor fluidly connected tothe compressor section, a turbine section fluidly connected to thecombustor and mechanically connected to the compressor section via ashaft, a plurality of rotors disposed in one of the compressor sectionand the turbine section, each of the rotors including a rotor diskportion including a radially inward bore, and each of the rotors beingstatic relative to the shaft, each rotor in the plurality of rotorsbeing axially adjacent at least one other of the rotors in the pluralityof rotors and defining a gap between each of the rotors and the axiallyadjacent rotors, a cooling passage for a cooling flow defined betweenthe shaft and the rotors, and a cooling flow redirection componentdisposed at the gap and operable to redirect the cooling flow in thecooling passage into the gap.

In a further example of the above embodiment, the gap is defined betweenradially aligned surfaces of adjacent rotor disks.

In a further example of any of the above embodiments, the cooling flowredirection component includes a radially outward protrusion from theshaft.

In a further example of any of the above embodiments, cooling flowthrough the cooling passage contacts a planar surface of the coolingflow redirection component.

In a further example of any of the above embodiments, cooling flowthrough the cooling passage contacts a curved surface of the coolingflow redirection component.

In a further example of any of the above embodiments, the cooling flowredirection component comprises a cooling flow redirection portion, thecooling flow redirection portion being angled relative to flow throughthe cooling passage such that coolant flow is redirected into the gap.

In a further example of any of the above embodiments, the cooling flowredirection component further comprises a radially aligned portionextending radially outward from the cooling flow redirection portion,relative to a radius of the engine, and wherein the radially alignedportion includes at least one through hole.

In a further example of any of the above embodiments, the through holeis a tapered through hole having an inlet with a first cross sectionalarea, and an outlet with a second cross sectional area, and wherein thesecond cross sectional area is smaller than the first cross sectionalarea.

In a further example of any of the above embodiments, the cooling flowredirection component is integral to one of the rotors defining the gap.

A further example of any of the above embodiments includes a secondarycooling flow inlet positioned at a radially outward edge of the gap.

In a further example of any of the above embodiments, the cooling flowredirection feature includes at least one cooling flow accelerationcomponent positioned on an upstream surface of the cooling flowredirection feature, relative to cooling flow through the cooling flowpassage.

In another example embodiment of the following disclosure, a method forcooling rotor bores includes providing a cooling flow path radiallyinward of a set of rotor bores and radially outward of a shaft in one ofa turbine section and a rotor section of a turbomachine, redirecting atleast a portion of cooling flow in the cooling flow path into a gapdefined between adjacent rotor bores, thereby increasing cooling of therotor bores, and returning at least a portion of the redirected coolingflow to the cooling flow path.

In a further example of any of the above embodiments, redirecting atleast a portion of cooling flow in the cooling flow path into a gapdefined between adjacent rotor bores, thereby increasing cooling of therotor bores comprises disposing a cooling flow redirection component inthe cooling flow path at the gap.

In a further example of any of the above embodiments, disposing acooling flow redirection component in the cooling flow path at the gapcomprises disposing a first surface of the cooling flow redirectioncomponent in the flow path, such that a cooling flow contacting thefirst surface is redirected into the gap.

In a further example of any of the above embodiments, the first surfaceis approximately normal to the cooling flow.

In a further example of any of the above embodiments, the first surfaceis a curved surface.

A further example embodiment of any of the above embodiments includessupporting the cooling flow redirection component on the shaft at thegap.

A further example embodiment of any of the above embodiments includessupporting the cooling flow redirection component on at least one of setof rotor bores.

In one example embodiment of the following disclosure, a cooling flowredirection component includes a first cooling flow redirection surfaceoperable to interfere with a cooling flow and redirect the cooling flowradially outward, and an interconnection feature operable tointerconnect the cooling flow redirection component with at least one ofa first rotor bore defining a gap, a second rotor bore defining a gapand a shaft defining a cooling flow passage radially inward of the rotorbores.

A further example embodiment of any of the above embodiments includes ahole in the cooling flow redirection component, wherein the hole directscooling flow passing through the hole such that the cooling flow passingthrough the hole forms an impingement cooling jet.

These and other features of the present invention can be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically illustrates an example gas turbine engine.

FIG. 2 schematically illustrates a partial view of a turbine section forthe gas turbine engine of FIG. 1.

FIG. 3 schematically illustrates a first example cooling featuredisposed in a gap defined by two adjacent rotor disks.

FIG. 4 schematically illustrates a second example cooling featuredisposed in a gap defined by two adjacent rotor disks.

FIG. 5 schematically illustrates a third example cooling featuredisposed in a gap defined by two adjacent rotor disks.

FIG. 6 schematically illustrates a fourth example cooling featuredisposed in a gap defined by two adjacent rotor disks.

FIG. 7 schematically illustrates a fifth example cooling featuredisposed in a gap defined by two adjacent rotor disks.

DETAILED DESCRIPTION OF AN EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

Within each of the compressor section 24 and the turbine section 28 aremultiple rotors. Each of the rotors includes a rotor disk positionedabout the axis and connected to one of the shafts 40, 50. The connectionbetween each rotor disk and the corresponding shaft can be either adirect connection or an indirect connection. Protruding radially outwardfrom each rotor disk are multiple rotor blades. The rotor blades extendradially outward into the core flow path C. interspersed between eachrotor and the adjacent rotor or rotors in the compressor section 24 orturbine section 28 are multiple stators. The combination of the rotorsand the stators defines the flow path in the compressor section 24 orthe turbine section 28.

With continued reference to FIG. 1, FIG. 2 schematically illustrates apartial view of a rotor/stator arrangement 100 such as could be utilizedin either a compressor section 24 or a turbine section 28 of the gasturbine engine 20 of FIG. 1. A first rotor disk 110 includes a rotorbore 116 and a radially outward portion 114. Protruding from theradially outward portion 114 is a rotor blade 112. Immediatelydownstream of the first rotor 110, in the core flow path C, is a stator130. An inner diameter 132 of the stator 130 is interfaced with theradially outward portion 114 of the rotor disk 110 via a seal 134. Theseal 134 allows the rotor disk 110 to rotate relative to the stator 130and maintains the integrity of the core flow path C. Immediatelydownstream of the stator 130 is a second rotor disk 120 including arotor bore 126 and a radially outward region 124. As with the firstrotor disk 110, the second rotor disk 120 includes multiple rotor blades122 extending radially outward from the rotor disk 120 into the coreflow path C. The first rotor disk 110 and the second rotor disk 120 areconsidered to be adjacent to each other, as there are no interveningrotor disks 110, 120.

Defined between the rotor disks 110, 120 and a shaft 140 is a coolingpassage 150. Cooling air is passed through the cooling passage 150 alonga cooling flow path 152. The cooling air cools the rotor bores 116, 126as the air passes through the cooling passage 150.

The rotor disks 110, 120 are mechanically connected via a joint 180. Thejoint 180 ensures that the rotor disks 110, 120 rotate together. Thesecond rotor disk 120 further includes an extension 182 that isconnected to a further rotor connected to the shaft 140, or is connecteddirectly to the shaft 140. Due to the connection between the rotor disk120 and the shaft 140, the rotor disks 110, 120 and the shaft 140 spintogether, and the rotor disks 110, 120 are static relative to the shaft140.

Defined between the first and second rotor disks 110, 120 and radiallyinward of the joint 180 is a gap 170. Further defining the gap 170 is aradially aligned surface 172 of the first rotor disk 110 and a radiallyaligned surface 174 of the second rotor disk 120.

During operation of the gas turbine engine 20, the rotor blades 112, 122and the rotor disks 110, 120 are exposed to extreme temperatures, andextreme temperature changes. As the radially outward portions 114, 124of the rotor disks 110, 120 and the rotor blades 112, 122 aresignificantly thinner than the rotor bores 116, 126 of the rotor disks110, 120, the radially outward regions 114, 124 and the rotor blades112, 122 heat up and cool down significantly faster than the rotor bores116, 126. The different heating and cooling rates can result in largeradial thermal gradients on the rotor disks 110, 120. The large thermalgradients impart significant thermal stress on the rotor disks 110, 120and reduce the lifespan of the rotor disks 110, 120.

In order to reduce the thermal gradients, a cooling flow 152 is providedin the cooling passage 150. The cooling flow 152 contacts a radiallyinward surface of each of the rotor disks 110, 120 and providesadditional cooling to the rotor bore regions 116, 126. A minimal amountof the cooling flow 152, if any, leaks into the gap 170, and minimalcooling is achieved in the gap 170. In some examples, a heating flow canbe provided along the same path 152, when heating of the rotor bores116, 126 is desired.

With continued reference to FIG. 2, and with like numerals indicatinglike elements, FIG. 3 schematically illustrates adjacent rotor bores216, 226 defining a gap 270. As with the example of FIG. 2, the rotorbores 216, 226 and a shaft 240 define a cooling passage 250 throughwhich a cooling flow 252 flows. The two rotor bores 216, 226 are joinedat a joint 280, and are static relative to the shaft 240.

Disposed within the gap 270 is a cooling flow redirection component 290.In the example of FIG. 3, the cooling flow redirection component 290 isa radial protrusion from the shaft 240. The cooling flow redirectioncomponent 290 protrudes into the gap 270 and includes an upstream facingsurface 292. The upstream facing surface 292 intersects with the coolingflow 252 and redirects the cooling flow 252 radially outward from thecooling flow 252 into the gap 270. The illustrated example surface 292is approximately normal to the cooling flow 252 and causes theredirected cooling flow 254 to be turbulent as the redirected coolingflow 254 passes through the gap 270. The introduced turbulence furtherincreases the contact between the redirected cooling flow 254 and theradially aligned surfaces defining the gap 270. The increased contact,in turn increases the magnitude of the cooling provided to thecorresponding rotor bore region 216, 226.

After passing through the gap 270, the redirected cooling flow 254returns to the cooling passage 250. In some examples, a flow correctingfeature 294, such as a flow guide, is included in the gap 270 to removeturbulence from the redirected cooling flow 254 as the flow returns tothe cooling passage 250.

The increased cooling provided by the cooling flow redirection component290 and the redirected cooling flow 254 increases the cooling rate ofthe rotor bores 216, 226, thereby decreasing the thermal gradients onthe rotor disks 110, 120 (illustrated in FIG. 2).

In some examples, the cooling flow redirection component 290 is integralto the shaft 240. In alternate examples, the cooling flow redirectioncomponent 290 is an additional component installed on the shaft 290 aspart of engine assembly.

With continued reference to FIGS. 2 and 3, and with like numeralsindicating like elements, FIG. 4 schematically illustrates secondexample configuration of adjacent rotor bores 316, 326 defining a gap370. As with the example of FIG. 2, the rotor bores 316, 326 and a shaft340 define a cooling passage 350 through which a cooling flow 352 flows.The two rotor bores 316, 326 are joined at a joint 380, and are staticrelative to the shaft 340.

As with the example of FIG. 3, disposed within the gap 370 is a coolingflow redirection component 390. In the example of FIG. 4, the coolingflow redirection component 390 is radial protrusion from the shaft 340.The cooling flow redirection component 390 protrudes into the gap 370and includes an upstream facing surface 392. The upstream facing surface392 is a curved surface that is concave relative to the flow directionof the cooling flow 352. The concave curved surface redirects thecooling flow 252 along a redirected cooling flow path 364 withoutintroducing turbulence. The redirected cooling flow path 364 travelsalong the radially aligned surfaces of the rotor bores 316, 326 definingthe gap 370. As the redirected cooling flow travels along the redirectedcooling flow path 354, the rotor bores 316, 326 receive increasedcooling. At the downstream edge of the gap 270, the redirected coolingflow 364 is returned to the cooling flow path 350.

Further, as with the example of FIG. 3, the cooling flow redirectioncomponent 390 can be constructed integral to the shaft 240 in someexamples, and can be a separate component installed onto the shaft aspart of the engine assembly in other examples.

With reference to both the cooling flow redirection component 290 ofFIG. 3, and the cooling flow redirection component 390 of FIG. 4, thecooling flow redirection components 290, 390 are elements that redirectthe cooling flow into the gap 270, 370 without restricting the coolingflow. Depending on the dimensions of the gap 270, 370, and thedimensions of the cooling flow redirection component 290, 390, coolingflow through the gap 270 can be slowed down, increasing exposure timebetween the cooing flow and the radially aligned surfaces of the boreregions 216, 316, 226, 326 or maintained the same.

With continued reference to FIGS. 2-4, and with like numerals indicatinglike elements, FIG. 5 schematically illustrates a third exampleconfiguration of adjacent rotor bores 416, 426 defining a gap 470. Aswith the example of FIG. 2, the rotor bores 416, 426 and a shaft 440define a cooling passage 450 through which a cooling flow 452 flows. Thetwo rotor bores 416, 426 are joined at a joint 480, and are staticrelative to the shaft 440.

Interconnecting in the joint 480, is a cooling flow redirectioncomponent 490. The cooling flow redirection component 490 has an angledsurface 492 on a cooling flow redirection portion 494. The angledsurface 492 is angled relative to the cooling flow 452 through thecooling passage 452. Protruding radially outward from the cooling flowredirection portion 494 is a radially aligned portion 496. The radiallyaligned portion 496 is interconnected in the joint 480. Further includedin the radially aligned portion 496 are one or more holes 498. Thecooling flow redirection component 490 splits the gap 470 into anupstream portion 470 a and a downstream portion 470 b, with the hole498, or holes, connecting the two portions 470 a, 470 b. In theillustrated example of FIG. 5, the walls of the hole 498 are alignedwith the axis of the engine 20, and do not further restrict flow betweenthe portions 470 a, 470 b.

During operation, the cooling flow 452 contacts the angled surface 492and is redirected radially outward as a redirected cooling flow 454. Theredirected cooling flow 454 passes along the radially aligned surface ofthe upstream rotor bore 416, cooling the upstream rotor bore 416. Theredirected cooling flow 454 then passes through the hole 498 and isdirected toward the radially aligned surface of the downstream rotorbore 426. The redirected cooling flow 454 impinges on the radiallyaligned surface of the downstream rotor bore 426 providing a furtherimpingement cooling effect. The redirected cooling flow 454 is thenreturned to the cooling flow path 452.

In the illustrated example of FIG. 5, the cooling flow redirectioncomponent 490 is a distinct component that is interconnected with thejoint 480 using any known interconnection technique during assembly ofthe engine 20. In alternate examples, the cooling flow redirectioncomponent 490 can be integrally formed with either the upstream rotordisk 110 (illustrated in FIG. 2) or the downstream rotor disk 120(illustrated in FIG. 2).

Turning our attention now to FIG. 6, and with continued reference toFIG. 5, FIG. 6 illustrates a fourth example configuration of adjacentrotor bores 516, 526 defining a gap 570. As with the example of FIG. 5,the rotor bores 516, 526 and a shaft 540 define a cooling passage 550through which a cooling flow 552 flows. The two rotor bores 516, 526 arejoined at a joint 480, and are static relative to the shaft 440.

The cooling flow redirection component 590 of FIG. 6, and the gap 570arrangement of FIG. 5 are similar to those of FIG. 5 with the exceptionof the hole 598. In the example of FIG. 6, the hole 598 is tapered suchthat the downstream opening of the hole 598 has a smaller crosssectional area than the upstream opening of the hole 598. As a result ofthe tapering, the cooling flow passing through the hole 498 isaccelerated, increasing the impingement on the radially aligned surfaceof the downstream rotor bore 526. Further, the specific angle of thetapering in the hole 598 can be adjusted and determined by one of skillin the art, having the benefit of this disclosure, to direct the coolingflow to a specific area of the radially aligned downstream bore andprovide increased cooling to the specific area.

With continued reference to each of the preceding examples, and withlike numerals indicating like elements, FIG. 7 illustrates the exampleof FIG. 4 with an additional cooling feature disposed in the joint 380.One of skill in the art, having the benefit of this disclosure willappreciate that the additional cooling feature in the joint 380 can beapplied to each of the preceding examples, and is not limited to theexample illustrated in FIG. 7.

The joint 380 in the example of FIG. 7 includes a passage 682,alternatively referred to as a cooling flow inlet, such as a throughhole. The passage 682 provides a secondary cooling flow path 684 throughthe joint 380 at a radially outward edge of the gap 370. The secondaryflow path 684 provides a supplemental cooling flow to increase thecooling flow through the redirected cooling flow 364 and the coolingflow path 352 downstream of the gap 370.

While each of the above examples is illustrated as a single cooling flowredirection component disposed in a gap between adjacent rotor disks,one of skill in the art, having the benefit of this disclosure willunderstand that in a practical gas turbine engine, multiple rotor diskswill be adjacent in a stack, resulting in the creation of multiple gaps.Cooling flow redirection components, as described above, can be includedin one or more of the gaps in a practical engine depending on thespecific cooling needs of the engine, and engines including multiplecooling flow redirection components are within the contemplation of thisdisclosure.

Further, while illustrated herein the context of a geared turbofanengine, one of skill in the art, having the benefit of this disclosurewill understand that the above cooling arrangements can be incorporatedinto any rotary machine, including direct drive turbines, land basedturbines, marine turbines, or the like, and is not limited to gearedturbine engines.

It is further understood that any of the above described concepts can beused alone or in combination with any or all of the other abovedescribed concepts. Although an embodiment of this invention has beendisclosed, a worker of ordinary skill in this art would recognize thatcertain modifications would come within the scope of this invention. Forthat reason, the following claims should be studied to determine thetrue scope and content of this invention.

The invention claimed is:
 1. A gas turbine engine comprising: acompressor section; a combustor fluidly connected to the compressorsection; a turbine section fluidly connected to the combustor andmechanically connected to said compressor section via a shaft; aplurality of rotors disposed in one of said compressor section and saidturbine section, each of said rotors including a rotor disk portionincluding a radially inward bore, and each of said rotors being staticrelative to said shaft; each rotor in said plurality of rotors beingaxially adjacent at least one other of said rotors in said plurality ofrotors and defining a gap between each of said rotors and said axiallyadjacent rotors, wherein each of said rotors includes a joint armpartially crossing said gap; a cooling passage for a cooling flowdefined between said shaft and said rotors; a cooling flow redirectioncomponent disposed at said gap and operable to redirect said coolingflow in said cooling passage into said gap, wherein the cooling flowredirection component is distinct from said rotors in said plurality ofrotors and includes a radially outward protrusion from said shaft, theradially outward protrusion including a curved surface.
 2. The gasturbine engine of claim 1, wherein said gap is defined between radiallyaligned surfaces of adjacent rotor disks.
 3. The gas turbine engine ofclaim 1, wherein cooling flow through said cooling passage contacts thecurved surface of said cooling flow redirection component.
 4. The gasturbine engine of claim 1, further comprising a secondary cooling flowinlet positioned at a radially outward edge of said gap.
 5. The gasturbine engine of claim 1, wherein the curved surface is concaverelative to a cooling flow direction.
 6. The gas turbine engine of claim1, wherein the curved surface is disposed in the cooling passage.